This invention relates to an axial-flow turbine blade, more particularly, to an axial-flow turbine rotor blade for gas turbine engines cooled by a cooling fluid such as compressor air flowing through several cooling fluid channels leading radially through the blade. Two cooling fluid channels communicate with each other through impingement cooling ducts, through which the cooling air fluid such as air entering the blade through the one channel is lead into the other channel nearer the leading edge of the blade for providing an impingement cooling. The cooling fluid is ejected from the other channel through lateral holes penetrating the wall of the blade at a slant or even tangentially.
German Patent Specification No. DE-AS 1,601,561 discloses such a turbine blade. In that turbine blade, a blade shell having a relatively thin wall completely encases several cooling channels which extend radially through the blade and communicate with each other through various ports or ducts. In this prior art blade the cooling air is admitted at the root end of the blade through a first radially extending channel and, from there, to a second air cooling channel through impingement air cooling ducts in a transverse wall, said second air channel extending essentially within the entire leading edge of the turbine blade. At the leading edge proper the blade shell has a relatively large number of radially and vertically staggered cooling air bleed holes for producing a cooling air film along the leading edge during operation. This arrangement provides a relatively good cooling of the leading edge of the blade, which during operation is exposed to high temperatures. However, the known structure does not exclude the risk that the minute leading edge cooling holes are blocked at least partially during long service. The holes can be blocked or clogged by, e.g., contaminations carried in the cooling medium or cooling air.
Another consideration in the manufacture of such blades is that making such small diameter holes e.g., by, electro erosive discharge machining, involves a considerable investment, especially if the required accurate inspection after manufacture is taken into account to make sure that these holes were actually completed to provide the specified flow cross-sectional areas.
It has also been shown that on such turbine blades, the impingement air cooling holes or ducts mentioned above are exposed to attack, e.g., by a hot gas corrosion, despite the described cooling provisions. Besides, the impingement cooling air holes or ducts are relatively difficult to inspect. Additionally, it is comparatively difficult to manufacture said impingement air cooling holes especially by electrochemical methods, because the need to guide the respective drilling tools through the air channels renders the drilling operation rather difficult and requires drilling tools of complex manufacture which additionally are difficult to handle.
Further, the known cooling methods for the blade require relatively high cooling air pressures and cooling air velocities involving considerable aerodynamic losses especially for the intended film cooling of the leading edge.